Over the years, gas turbine engine manufacturers have increased the temperature and pressure at which gas turbine engines operate to meet demands for more powerful and efficient engines. The increased temperature and pressure levels have imposed rigorous operating conditions on certain engine components, particularly turbine vanes and blades immediately downstream of a combustor. In modern engines, turbine vanes and blades may be exposed to temperatures above the melting point of the alloy from which they are made.
While manufacturers have been designing gas turbine engines that operate under very demanding conditions, they have been striving to improve gas turbine engine reliability and to extend maintenance intervals to improve the economics of operating gas turbine engines. Manufacturers have addressed both objectives by applying protective coatings to certain parts, particularly turbine vanes and blades. Initially, the coatings focused on providing oxidation and corrosion protection. Examples of these include overlay and diffusion aluminide coatings, MCrAIY coatings, where M is Ni, Co, Fe, or Ni/Co, and other metallic coatings. Commonly assigned U.S. Pat. Nos. 4,585,481 and Re 32,121, both to Gupta et al., describe such coatings. More recently, multi-layer, thermal barrier coatings (TBC) that comprise an oxidation and corrosion resistant metallic bond coat and a thermally insulating ceramic top coat have been used. Such coatings are described in commonly assigned U.S. Pat. No. 4,321,310 to Ulion et al., U.S. Pat No. 4,321,311 to Strangman, U.S. Pat No. 4,401,697 to Strangman, U.S. Pat No. 4,405,659 to Strangman, U.S. Pat No. 4,405,660 to Ulion et al., U.S. Pat No. 4,414,249 to Ulion et al., and U.S. Pat No. 5,262,245 to Ulion et al. Thermal barrier coatings provide thermal resistance to the high temperatures in a gas turbine engine in addition to providing oxidation and corrosion resistance.
For gas turbine applications, the materials and processing methods chosen for the thermal barrier coatings are selected to provide resistance to spallation (coating loss) of the ceramic outer layer during thermal cycling of the engine as well as resistance to the oxidizing and corrosive environment in the case of a TBC spallation event. During normal engine operation and after time, the thermal barrier coating, including the metallic bond coat and the ceramic top coat, will degrade in certain surface areas most subjected to strenuous operating conditions. The bond coat may interdiffuse with an article substrate in such surface areas during operation to the extent that its protective ability has been reduced below an acceptable level, requiring the removal and reapplication of a protective coating.
In addition, internal cooling techniques have been developed to keep the temperature of the vanes and blades within design limits while operating at high temperatures. For example, the outer surface of engine components exposed to the hot gas path are typically cooled with high pressure cooling air from the compressor section of the engine. Film cooling has proven to be an effective means of utilizing this cooling air. In this method, a layer of cool air is flowed between the high temperature gases and the external surfaces of the engine components. The layer of cooling air is formed by passing the cooling air through a series of small cooling holes in the component which are formed in a predetermined pattern. The resulting film of air reduces component surface temperature thereby deterring component distortion. Engine efficiency is also increased because higher turbine inlet temperature ranges are possible.
It is well known in the art that film cooling effectiveness can be increased by using diffusion holes that have a conical portion and an enlarged opening at the surface of the component. The shaping of the holes to diffuse air before it enters the boundary layer of the component broadens the spread of air downstream of the hole and thus, increases cooling effectiveness. In comparison, cylindrical shaped holes create a localized region downstream of the hole where cooling effectiveness decay is minimized. Although high quality diffusion holes provide superior performance, they are both costly and difficult to form.
Because turbine blades and vanes are expensive, a variety of refurbishment techniques have been developed to restore the deteriorated vanes to serviceable condition. The specific details of the various refurbishment techniques depend on the nature and extent of vane deterioration. For instance, existing protective coatings, such as, the thermal barrier coatings that include the bond coat and the ceramic top coat, may be removed from the blades and vanes.
Removal of the bond coat after removal of the ceramic top coat may be required due to surface degradation of the bond coat especially in those surface areas most subject to strenuous operating conditions. The ceramic portion of the coating may be stripped by soaking the part in a solution of KOH. The metallic portion of the coating may be stripped by soaking the part in a HCl solution.
Prior to reapplying a non-original replacement coating and after removal of the existing thermal barrier coating, a repair of cracks and other surface defects in the vane and blade castings may take place. Such a repair process is described in U.S. Pat. No. 4,008,844. According to this patent, a repair material comprises a mixture of metal powders made from two powders with different compositions. One composition approximates that of the superalloy to be repaired while the other composition also approximates the superalloy to be repaired, but contains a melting point depressant, usually boron. The mix has a paste-like consistency. The defect to be repaired is filled with a mixture of these powders and then heated to a temperature at which the boron containing the powder melts, but the boron-free powder and the substrate do not. Solidification then occurs isothermally over a period of time as the boron diffuses into the substrate thereby raising the solidification temperature of the melted constituent. Typically, all the cooling holes, for example in the vane, which depending of the airfoil can be in excess of about 300, are completely filled with the repair material. The filing process is both labor intensive and costly and will necessitate the remanufacture of the filled cooling holes, including the diffusion holes.
As is known in the case of blade repair, the blade may first be stripped of its original coating and then a nonoriginal replacement coating is applied to the blade casting prior to returning the blade to service. During this repair process, if the blade should have any cooling holes, these cooling holes may be subject to being partially or completely filled with the non-original coating material.
Such excess non-original coating can accumulate in the mouth of each cooling hole. This phenomenon is known as “coatdown” and can restrict the flow capacity of the affected holes. The effects of coatdown can diminish the cooling effectiveness of the film cooling thereby reducing the component's useful operating life. Any cooling holes that are subject to coatdown are typically unacceptable for return to service and will require reworking to remove the excess nonoriginal coating before the blade can be put back into service.
The effects of coatdown can be reversed by eroding the excess coating by propelling a high velocity, precisely focused stream of abrasive particles into the mouth of each affected hole. However, the erosive treatment can be inaccurate and nonrepeatable and is tedious and time consuming since a typical turbine airfoil has many rows of cooling holes.
Therefore, the repair of turbine components require the remanufacture of the cooling holes typically employing the processes used in the original manufacture of the component.
Many attempts have been made to remanufacture cost effective, high quality cooling holes in gas turbine engine components. For example, laser drilling has been used to produce cylindrical holes on the leading and trailing edges of vanes and blades. It is difficult, however, to produce shaped holes (diffusion holes) with this technique. This is a significant repair limitation because the geometry of the holes partially determines the effectiveness of cooling.
Electrical discharge machining (EDM) is a well-known process for producing shaped holes or other openings in metals. It uses current discharges to erode metal. For example, by pulsing a direct current between a positively charged work piece (anode) and an electrode (cathode), a spark discharge may be produced. The current occurs when the potential difference between the electrode and the work piece, which both contact a dielectric fluid, is great enough to breakdown the dielectric fluid and produce an electrically conductive channel. Upon application of a voltage or potential, a current flow results with enough heat energy to melt and erode the work piece. This process has application in the machining of small, deep, odd-shaped holes which are cumbersome to produce by other means.
An EDM method for producing or remanufacturing diffusion holes in engine components uses a copper electrode that is manufactured in a three-dimensional shape by stamping and coining. The electrode consists of at least one small diameter elongated end that produces the cooling air metering section. The elongated end is connected to a three-dimensional diffuser shaped portion that produces a diffuser area for the metering section. The electrode produces a similar shaped hole, with allowance for electrode overburn and EDM electrode erosion. Although the above EDM method is successful, limitations exist. EDM is a time intensive and relatively expensive process compared to other processes such as laser drilling. Also, the electrodes are fragile and are not reusable. The use of EDM to remanufacture the diffusion cooling holes in a typical vane is labor intensive and expensive.
Thus, what is needed in the gas turbine industry is a repair for gas turbine engine components, and in particular vanes, having diffusion cooling passages that permits removal of the entire coating system and the repair of defects while increasing the number of times a part can be repaired and reducing manufacturing cost and cycle time, as compared to prior art repair methods.